A conventional gas turbine engine includes a compressor, a combustor and a turbine. The compressor compresses ambient air which is supplied to the combustor where the compressed air is combined with a fuel and ignites the mixture, creating combustion products defining a working gas. The working gas is supplied to the turbine where the gas passes through a plurality of paired rows of stationary vanes and rotating blades. The rotating blades are coupled to a shaft and disc assembly. As the working gas expands through the turbine, the working gas causes the blades, and therefore the shaft and disc assembly, to rotate.
Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine blades must be made of materials capable of withstanding such high temperatures. In addition, turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures.
Typically, turbine blades comprise a root, a platform and an airfoil that extends outwardly from the platform. The airfoil is ordinarily composed of a tip, a leading edge and a trailing edge. Most blades typically contain internal cooling channels forming a cooling system. The cooling channels in the blades may receive air from the compressor of the turbine engine and pass the air through the blade. The cooling channels often include multiple flow paths that are designed to maintain the turbine blade at a relatively uniform temperature. However, centrifugal forces and air flow at boundary layers often prevent some areas of the turbine blade from being adequately cooled, which results in the formation of localized hot spots. Localized hot spots, depending on their location, can reduce the useful life of a turbine blade and can damage a turbine blade to an extent necessitating replacement of the blade.
A conventional cooling system in a turbine blade assembly may include an intricate maze of cooling flow paths through various portions of the turbine blade.
Further, third row turbine blades, which may comprise a highly tapered large airfoil, present additional cooling problems associated with the geometry of the airfoil. Specifically, the lower or radially inward portion of the airfoil comprises a large cross section area, and the cross section tapers to a smaller thickness toward the tip of the blade. Accordingly, the configuration of the cooling circuit requires particular consideration to providing an airflow in contact with the lower portions of the airfoil to maintain the heat transfer coefficient in the wider cross section portion of the blade.
While many of the known cooling systems for turbine blades have operated successfully, a need still exists to provide increased cooling capability, particularly in turbine blades having highly tapered large airfoils.